Developer Notes

The purpose of this document is to explain the inner workings of MachUpX for those who are developing it. Please update this as often as you can.

Numerical Lifting-Line Implementation

MachUpX solves the formulation of the numerical lifting-line equation presented in Cory Goates’ Master’s thesis. This formulation differs somewhat from the original formulation by Phillips and Snyder. Most significantly, the formulation is generalized to allow multiple aircraft to be analyzed at once. MachUpX sets up and solves the lifting-line equation in Earth-fixed coordinates. Once the distribution of vortex strengths is obtained, the forces and moments are integrated then transformed to the body-fixed frame for each aircraft.

Multiple Aircraft

For the most part, MachUpX is ignorant of how many aircraft are in the scene. When performing calculations, it will simply execute a for loop on the stored dictionary of aircraft.

API to Aircraft

When using MachUpX as a Python module, the Airplane class is never exposed to the user. Rather, the user calls getters and setters through the Scene class and supplies the name of the target aircraft to effect changes.

In the future, it could be beneficial to allow the user access to the Airplane class directly. The user could instantiate and aircraft object and then pass this object as an argument to the Scene constructor. This would require some kind of listener/binding between the Airplane class and the Scene class so that certain functions in the Scene class, such as _perform_geometry_calculations(), are called whenever a change is made to the Airplane object. This is just an idea.

Wing Segment Tree Structure

In implementing the WingSegment class, the set of wing segments constituting an airplane are stored as a tree. There is an origin segment, which has no properties. Other segments are then attached to (or more accurately, defined relative to) this segment. Each segment knows which segments are attached to it and segments are added recursively. This structural choice was motivated by two factors. For one, recursion is elegant, simple, and fast. The second is that this will more naturally allow for calculating the influence of forces and moments exerted by outboard wing segments on inboard segments. The tree can be traversed recursively and forces, moments, and aeroelastic deflections can be integrated as the function backs out of the recursion. Aeroelastic effects are not yet implemented in MachUpX and so this may change in the future as needed functionality becomes clearer.

Mean Aerodynamic Chord (MAC) Calculation

MAC seems to be a poor man’s estimate for the aerodynamic center, but not quite. It is based entirely off of geometry, no aerodynamics involved. I took information from the following to code this up:

  • Raymer “Simplified Aircraft Design for Homebuilders”
  • McCormick “Aerodynamics, Aeronautics, and Flight Mechanics”
  • Nickel & Wohlfahrt “Tailless Aircraft in Theory and Practice”
  • Phillips “Mechanics of Flight”

There is agreement between all sources on only one thing, the definition of the MAC as a reference length. This is given by Phillips Eq. 1.8.86 (excuse my ASCII math):

            b/2
      1    /     2
MAC = -    |    c  dy
      S    /
       -b/2

Phillips and Raymer are the only ones to tie the MAC to a specific location on the wing. Phillips simply states that there is a spanwise position on the wing where the local chord is equal to the MAC. Raymer also does this and presents a graphical method for determining the position of the MAC on a tapered wing. He briefly describes how to find the MAC for a compound tapered wing, stating the MAC should be found for each section separately and the the 25% point should be averaged using an area-weighted average. N&W and McCormick simply give the MAC as a reference length and give it no positional significance.

Raymer ties the MAC to an introductory discussion of stability. He states, “The MAC is a sort of averaged chord, and the entire wing tends to act as if all its area were concentrated at the MAC.” He thereby claims the point of neutral stability (i.e. the aerodynamic center) on the wing is at the 25% MAC location. He then recommends, for a tailed airplane, that the CG be placed at the 25% MAC location so the wing is neutrally stable and all longitudinal stability comes from the horizontal stabilizer. This is reasonable, given his assumptions.

Phillips directly disputes the assumption that the wing aerodynamic center lies on the MAC. He shows this is not true for most wing geometries. However, he focuses that the aerodynamic center does not lie on the same spanwise position as the MAC. For unswept wing, Phillips states the aerodynamic center of the wing lies on the locus of section aerodynamic centers (i.e. 25% root chord). This agrees with Raymer’s claim, as the 25% MAC location also lies on the straight locus of section aerodynamic centers. For swept wings, determining the location of the wing aerodynamic center becomes more hairy. He gives the following approximation (Eq. 1.11.9):

                /     b/2            \
_      1    d  |     /       ~       |
x   = ----  -- |     |  CL c x   dy  |
 ac   CL,a  da |     /        ac     |
                \-b/2               /

If we assume the section lift coefficient is constant over the span, as is often done, this reduces to:

            b/2
_     1    /     ~
x   = -    |   c x   dy
 ac   S    /      ac
       -b/2

This can be interpreted as an area-weighted average of the section aerodynamic centers. N&W call this the “C-Point” or “Center of Lift for Constant Local Lift Coefficient”. This fits. They then recommend the CG be placed at least 6-12% of the MAC in front of the C-Point for a flying wing. As the main wing constitutes the entire airframe of a flying wing, this is simply the recommendation that the aircraft have a static margin of at least 6-12%. Very reasonable.

I have coded up both methods of determining the main wing aerodynamic center. For unswept wings, the 25% MAC location and the average of the section aerodynamic centers are the same, i.e. 0. For swept straight and tapered wings, they are also the same. For swept elliptic wings, however, the two results diverge.